Advanced booster system

ABSTRACT

A compression stage having a plurality of stator vanes and rotor blades coaxial with a longitudinal centerline axis, each stator vane having an exit swirl angle distribution such that the exit swirl angle has a maximum value at an intermediate radius location and each rotor blade having a blade leading edge adapted to receive the flow from the stator vanes with the exit swirl angle distribution profile.

CROSS REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. Ser. No. 11/606,759, filed onNov. 30, 2006, the entire disclosure of which is incorporated herein byreference.

BACKGROUND OF THE INVENTION

This invention relates generally to gas turbine engines, and, morespecifically, to the compression modules therein, such as the boosterand the compressor.

In a turbofan aircraft gas turbine engine, air is pressurized in a fanmodule and a compression module during operation. The air passingthrough the fan module is used for generating the bulk of the thrustneeded for propelling an aircraft in flight. The air channeled throughthe compression module is mixed with fuel in a combustor and ignited,generating hot combustion gases which flow through turbine stages thatextract energy therefrom for powering the fan and compressor rotors.

A typical compression module in a turbofan engine includes a multi stagebooster which compresses the air to an intermediate pressure and passesit to a multistage axial flow compressor which further pressurizes theair sequentially to produce high pressure air for combustion. Both thebooster and the compressor have rotor stages and stator stages. Thebooster rotor is typically driven by a low pressure turbine and thecompressor rotor is driven by a high pressure turbine.

Fundamental in booster and compressor design is efficiency incompressing the air with sufficient stall margin over the entire flightenvelope of operation from takeoff, cruise, and landing. However,compressor efficiency and stall margin are normally inversely relatedwith increasing efficiency typically corresponding with a decrease install margin. The conflicting requirements of stall margin andefficiency are particularly demanding in high performance jet enginesthat require increased power extraction, while still requiring high alevel of stall margin in conjunction with high compressor efficiency. Inconventional designs, efficiency is usually sacrificed in order toachieve improved operability and increased stall margin.

Operability of a compression system in a gas turbine engine istraditionally represented on an operating map with inlet corrected flowrate along the X-axis and the pressure ratio on the Y-axis, such as forexample, shown in FIG. 1 for a booster. In FIG. 1, operating line 102and the stall line 101 are shown, along with several constant speedlines 104-108. Line 104 represents a lower speed line and line 105represents a higher speed line as compared to the design speed line 103.As the booster is throttled from the operating line 102 at a constantspeed, such as the design speed represented by the constant speed line103, the inlet corrected flow rate decreases while the pressure ratioincreases, and the booster operation moves closer to the stall line 101.In order to avoid a stall, the fans, boosters and compressors in a gasturbine engine are designed to have sufficient stall margin with respectto the stall line, such as line 101 shown in FIG. 1.

Maximizing efficiency of booster and compressor airfoils is primarilyeffected by optimizing the velocity distributions over the pressure andsuction sides of the airfoil. However, efficiency is typically limitedin conventional booster and compressor designs by the requirement for asuitable stall margin. Any further increase in efficiency results in areduction in stall margin, and, conversely, further increase in stallmargin results in decrease in efficiency.

High efficiency is typically obtained by minimizing the wetted surfacearea of the airfoils for a given stage to correspondingly reduce airfoildrag. This is typically achieved by reducing airfoil solidity or thedensity of airfoils around the circumference of a rotor disk, or byincreasing airfoil aspect ratio of the chord to span lengths.

For a given rotor speed, this increase in efficiency reduces stallmargin. To achieve high levels of stall margin, a higher than optimumlevel of solidity may be used, along with designing the airfoils atbelow optimum incidence angles. This reduces axial flow compressorefficiency.

Increased stall margin may also be obtained by increasing rotor speed,but this in turn reduces efficiency by increasing the airfoil Machnumbers, which increases airfoil drag. Obtaining adequate stall marginis a problem especially in the case of the booster. Boosters typicallyare run at relatively lower wheel-speeds, while at the same time, thethroughflow velocity of the air is high. The booster is also unique ingeometry because the air flowing through the rear stages of the boosteris subjected to a significant change in direction of flow radiallyinward towards the longitudinal centerline axis. This results in aradial incidence swing imbalance as the booster is throttled to stallwith large incidence swings in the hub region of the airfoils. In thebooster, across the cruise and high power operating range where thebooster bleed valve is closed, stall typically initiates in the hubregion first, and therefore the incidence swings in the hub region areparticularly detrimental to operability. The incidence swings in the hubregion and the resulting stall margin loss become even more severeduring engine operation when there is increased demand for auxiliaryelectric power from the high pressure spool in the engine. Inconventional designs, efficiency is typically compromised to meetoperability requirements.

Embodiments are therefore provided that further improve the stall marginof the boosters and other high through-flow/wheel-speed compressorswithout significantly sacrificing the efficiency for improving gasturbine engine booster and compressor performance.

BRIEF DESCRIPTION OF THE INVENTION

A compression stage having a plurality of stator vanes and rotor bladescoaxial with a longitudinal centerline axis, each stator vane having anexit swirl angle distribution such that the exit swirl angle has amaximum value at an intermediate radius location and each rotor bladehaving a blade leading edge adapted to receive the flow from the statorvanes with the exit swirl angle distribution profile.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention, in accordance with preferred and exemplary embodiments,together with further objects and advantages thereof, is moreparticularly described in the following detailed description taken inconjunction with the accompanying drawings in which:

FIG. 1 is an example of the operating map of a booster, showingoperating line, stall line and the speed lines.

FIG. 2 is an axial sectional view through a portion of a gas turbineengine fan and booster.

FIG. 3 is an axial sectional view through a booster including rotorstages disposed axially between corresponding stator stages inaccordance with an exemplary embodiment of the present invention.

FIG. 4 is an axial view of a part of the booster rotor and stator stagesshowing a stator vane and corresponding rotor blades.

FIG. 5 is a radial sectional view through the airfoil of one of thestator vanes in a booster.

FIG. 6 is a comparison of an exemplary exit swirl angle distribution fora stator vane in accordance with an exemplary embodiment of the presentinvention with a conventional exit swirl angle distribution.

FIG. 7 is a plot of a set of exemplary exit swirl angle distributions,in normalized form, for the various stages of an exemplary boostersystem.

FIG. 8 is an exemplary embodiment of stator leading edge sweep anglevariations with span height for multiple stator stages of a booster.

FIG. 9 is an exemplary embodiment of rotor leading edge sweep anglevariations with span height for multiple rotor stages of a booster.

FIG. 10 is a radial sectional view through the airfoil of one of therotor blades in a booster, showing the location of the maximum airfoilthickness.

FIG. 11 is an exemplary distribution of the location of maximum airfoilthickness for airfoil sections at various span heights.

FIG. 12 is an exemplary embodiment of rotor trailing edge dihedral anglevariations with span height for multiple rotor stages of a booster.

DETAILED DESCRIPTION OF THE INVENTION

While the invention has been described in terms of various specificembodiments, those skilled in the art will recognize that the inventioncan be practiced with modification within the spirit and scope of theclaims.

Illustrated in FIG. 2 is a portion of a gas turbine engine fan 5 andbooster 7 configured for channeling and pressurizing a bypass airflow 2and a core airflow 3 respectively. The booster 7, which pressurizes theair flowing through the core, is axisymmetrical about a longitudinalcenterline axis 15, and includes an inlet guide vane (IGV) stage 11having a plurality of inlet guide vanes 12 spaced in a circumferentialdirection around the longitudinal centerline axis 15, a plurality ofstator vane stages 17. The booster 7 further includes multiple rotorstages 18 which have corresponding rotor blades 50 extending radiallyoutwardly from a rotor hub 19 or corresponding rotors in the form ofseparate disks, or integral blisks, or annular drums in any conventionalmanner.

Cooperating with each rotor stage, such as for example, the rotor stage18, is a corresponding stator stage 17. Each stator stage 17 in thebooster 7 comprises a plurality of circumferentially spaced apart statorvanes 40. The arrangement of stator vanes and rotor blades is shown inFIG. 4. The rotor blades 50 and stator vanes 40 define airfoils havingcorresponding aerodynamic profiles or contours for pressurizing the coreair flow 3 successively in axial stages. In operation, pressure of theair is increased as the air decelerates and diffuses through the statorand rotor airfoils.

FIG. 5 shows an exemplary radial sectional of the stator vane airfoil ina two dimensional axial plane view. As shown in FIG. 5, each stator vane40 defines an airfoil including a generally concave pressure side 44 anda circumferentially opposite, generally convex suction side 45. The twosides 44,45 extend chordally between an upstream leading edge 42 and anaxially opposite, downstream trailing edge 43. The booster is a high“throughflow-velocity/wheel speed” design, wherein it is driven by lowpressure turbines with relatively lower speeds, while the axial air flowvelocity through the booster is relatively high. Additionally, the hubflow pathway though the booster turns radially inward towards the enginecenterline. This causes the radial incidence angle to the airfoils toundergo large variations, especially in the hub region, as the boosteroperates in various flight regimes with varying demands on airflow. Thisis undesirable because stall in a booster may typically originate nearthe hub region of the airfoils. In conventional designs, in order toachieve operability goals in the presence of the high radial incidenceangle swing imbalance, efficiency is typically sacrificed. It isdesirable to have a booster design where the requirements for the stallmargin, including auxiliary electric power extraction, can be achievedwithout sacrificing the efficiency.

One way of accomplishing this is by utilizing stator vanes 40 and rotorblades 50 designed to reduce incidence angle swings in the hub regionsof the booster system during operation. Incidence angle for a rotorblade is defined as the difference between the relative inlet air angle306 measured from the meridional direction (β1, see FIG. 10) and theinlet metal angle 305 determined by the camber line angle at the leadingedge measured from the meridional direction (β1*, see FIG. 10). “Deltaincidence” (ΔINCIDENCE) is the difference between the incidence angle atstall line 101 and the incidence angle on the operating line 102. Forstator vanes the same definitions for incidence angle and “Deltaincidence” apply, except that the air angle is measured from themeridional direction in the absolute frame of reference. An exemplarystator vane 40 reduces the incidence flow swing in the booster hubregion by using a trailing edge 43 having a particular exit swirl angleprofile. An exemplary exit swirl angle distribution 144 for theexemplary stator vane 40 is shown in FIG. 6. FIG. 6 is a plot of theexit swirl angle versus the percent-span. The incidence angle swing inhub region of the rotor blades and stator vanes of the booster isreduced by adopting a trailing edge 43 with a particular distributionfor the exit swirl angle 140 from the root 46 to the tip 48, where theexit swirl angle is defined as the air angle leaving the stator trailingedge measured from the meridional direction omitting any secondary floweffects (shown in a 2D axial plane view in FIG. 5). Conventional designstator vanes typically result in an approximately linear andmonotonically increasing swirl angle distribution, such as thedistribution 142 in FIG. 6, in the exemplary design of the stator vane40 shown in FIG. 4, the vane has a tailored exit swirl angledistribution profile such as, for example item 144 shown in FIG. 6, fromthe root 46 to the tip 48 of the stator vane 40 such that the exit swirlangle 140 has a maximum value at an intermediate radius location 148between a first radius location 146 and the tip 48.

In an embodiment of the exemplary stator vane 40, the maximum value forthe exit swirl angle (about 22 degrees) in the trailing edge 43 occursat a span location of about 70% span height from the root, with thelowest value of the exit swirl angle (about 7 degrees) occurring at theroot 46 of the trailing edge 43 and the tip 48 has an exit swirl angle(about 18 degrees) in between the root value and the peak value. Theincidence swing near the hub region of the booster is significantlyreduced as compared to a conventional vane resulting in increased stallmargin and improved efficiency for the booster.

Stall margins for different rotor/stator stages can be improved bysuitably designing the stator vane airfoils with trailing edge exitswirl angle distributions similar to the one shown in FIG. 6 item 144.The location of the peak value of trailing edge exit swirl angle 140could be chosen to be at 50% span or higher, preferably in the 60% to80% span range, with the lowest value occurring near the root 46 of thestator vane 40. The trailing edge exit swirl angle distributions for thevarious stator stages of a preferred embodiment of a booster system areshown in FIG. 7 on a non-dimensional basis, where the exit swirl angleat the tip 48 has been reduced to a level that is in the range of 65% to85% of the exit swirl angle difference between the maximum value alongthe span and the minimum value at the root 46.

In another embodiment of the new stator vane 40 described above, theleading edge 42 of the stator vane 40 is designed with a sweep angleprofile. Aerodynamic sweep is a conventional parameter represented by alocal sweep angle which is a function of the direction of the incomingair and the orientation of the airfoil surface in both the axial, andcircumferential or tangential directions. The sweep angle is defined indetail in the U.S. Pat. No. 5,167,489, and is incorporated herein byreference. In the sign convention used herein, the aerodynamic sweepangle is represented as a negative value (−) for forward sweep, and apositive value (+) for aft sweep. In another embodiment of the statorvane 40 with tailored exit swirl angle distribution as describedpreviously, the stator vane leading edge 42 is designed with a forwardsweep near the root 46 of the airfoil in the hub region of the booster.This combination of a stator vane leading edge 42 with a forward sweepnear the root of the airfoil in the hub region of the booster and atrailing edge 43 with specific trailing edge exit swirl angledistribution further improves the aerodynamic performance andoperability of the booster.

FIG. 8 shows exemplary stator vane leading edge sweep angledistributions along the span for the various stator stages of anexemplary multistage booster. In the preferred embodiment for a multistage booster, the sweep angle is negative between the root 46 and afirst span location 147 in FIG. 4 and is positive from the first spanlocation 147 to the tip 48. The span height from the root 46 at whichthe sweep angle changes from negative to positive (denoted by “H” inFIG. 8) in a stator vane 40 is a function of the axial location of theparticular stator vane stage. As the air travels axially within thebooster from the entrance to the exit, it has to undergo sharp turnstowards the longitudinal centerline axis 15 of the booster prior toentry into a compressor located downstream. In the exemplary embodimentof a booster system 7, the stator vane leading edge sweep angledistributions, as shown in FIG. 8, are such that the span height fromthe root 46 at which the sweep angle changes from negative to positiveis higher for stator stages located further aft in the booster system.It is possible that one or more of the stator stages at the all end ofthe booster may have stator vanes with leading edges that have a forwardsweep only along the entire span. In FIG. 8, for example, the statorstage denoted by “S5” is such a stage.

In an embodiment of the booster system 7, the span location from theroot 46 at which the leading edge sweep angle changes from negative topositive is about 25% for a forward stage (denoted by “S2” in FIG. 8),50% for an intermediate stage (denoted by “S3” in FIG. 8) and 70% for arear stage (denoted by “S3” in FIG. 8) while the aft-most stage (denotedby “S5” in FIG. 8) has no leading edge aft sweep. In the preferredembodiment of the booster system 7, all the stator stages have statorvanes 40 such that the leading edge forward sweep at the root 46 for astator vane 40 is larger for stator stages located further aft in thebooster system and the stator vanes 40 have tips 48 having less leadingedge forward sweep, or more aft sweep, than at the root 46. In thepreferred embodiment of the booster system 7, the stator vane leadingedge sweep angle at the root 46 is about −3 degrees for the forward-moststage, about −5 degrees for the next stage aft, about −15 degrees forthe rear stage and about −20 degrees for the rear-most stage. The statorvane 40 leading edge 42 sweep angle at the tip 48 is about 13 degreesfor the forward-most stage, about 7 degrees for the next stage aft,about 5 degrees for the rear stage and about −2 degrees for therear-most stage.

As illustrated in FIG. 2, the booster system 7 in a gas turbine enginecomprises multiple rotor stages 18, with each rotor stage havingmultiple rotor blades. These rotor blades for the various rotor stagesare shown in FIG. 3, for example, as item 10 for a stage 2 rotor, item30 for a stage 3 rotor, item 50 for a stage 4 rotor, and item 70, for astage 5 rotor. As shown in FIG. 3, the first booster rotor stage (markedas “R2”) is located immediately aft of the inlet guide vane stage(marked as “IGV”). Each of the other rotor stages, R3-R5, is associatedwith the stator stages axially forward and aft from it, with each statorstage having multiple stator vanes. These stator vanes for the variousstator stages are shown in FIG. 3, for example, as item 20 for statorstage 2, item 40 for stator stage 3, item 60 for stator stage 4 and item80 for stator stage 5. Air exiting from a stator stage enters thedownstream adjacent rotor stage and is further compressed by the rotorblades in the rotor stage. As described in detail before, the statorvanes in a stator stage are designed to have specific trailing edge andleading edge characteristics to improve the operability and efficiencyof the booster. The operability and efficiency are also influenced bythe mechanical and aerodynamic design of the rotor blades in thebooster. Stall margins and efficiency of a compression stage and thebooster system can be enhanced by adopting the specific designcharacteristics for the rotor blades as disclosed and described herein.

The reduced incidence swing in the hubs of the airfoils results in asteeper speedline shape. Such steeper speedlines are shown in FIG. 1(items 106, 107 and 108).

Blade sweep has been used in fan and compressor blade designs forvarious reasons such as noise reduction and performance improvement. Inone embodiment of the present invention of a new rotor blade 50, theblade leading edge 52 has a new sweep profile such that in the rate ofchange of leading edge sweep angle with respect to the span height has asubstantially constant value along most of the leading edge span. Inanother embodiment, the leading edge sweep angle has a first rate ofchange with respect to the span height that is substantially constantnear the blade root 54, in a blade inner span region 155, and has asecond rate of change respect to span height that is substantiallyconstant along the span up to the blade tip 55 in a blade outer spanregion 156. In an embodiment of the blade, the blade inner span region155 covers a span of about 10% span height measured from the blade root54. In another embodiment, the rate of change of the leading edge sweepangle with respect to the span height is substantially constant alongthe entire blade leading edge 52.

FIG. 9 shows an embodiment having an exemplary variation of the leadingedge sweep angle along the span. As shown in FIG. 9, the blade leadingedge 52 has a forward sweep (negative sweep angle) near the root of theblade and an aft sweep (positive sweep angle) away from the root region.The rate of change of the leading edge sweep angle with respect to spanheight and the location of the blade first height 151 on the bladeleading edge 52 where the transition from forward sweep to aft sweepoccurs are chosen such that the flow coming out of the stator vanes,such as for example, stator vane 40 in FIG. 4, enters the rotor blades,such as for example, blade 50 in FIG. 4, with increased efficiency andis directed towards the hub region of the rotor in a manner to increasethe operability and efficiency of the rotor. As discussed previously,stall in a booster typically originates near the hub region over thehigher power ranges where the booster bleed valve operates closed.Having the unique characteristics of the blade leading edge 52 describedherein increases the stall margin for the booster. In an embodiment ofthe booster, all the rotor stages have rotor blades that havesubstantially the same characteristic linear variation of the leadingedge sweep angle with span height, as shown in FIG. 9.

FIG. 10 shows a radial sectional view through the airfoil of anexemplary rotor blade. In another embodiment, the locations of themaximum thickness 302 (identified as “Tmax”, see FIG. 10) of the rotorblade airfoil sections 300 are chosen such that they are located closerto the leading edge 52 at higher span locations from the blade root 54and the relative distance of the Tmax location from the leading edgevaries in a substantially linear manner with respect to the span heightfrom the blade root 54 to the blade tip 55. In this context, the“relative distance” is defined as the ratio of the axial distance “d”303 (see FIG. 10) of the Tmax location along an axial line from theblade leading edge 52 to the axial chord length “C” 301 (see FIG. 10) ofthe airfoil section 300 at a particular span height.

Locating Tmax 302 near the blade leading edge 52 at higher span heightsfrom the blade root 54 results in higher wedge angles for the bladeleading edge 52 in the radially outer sections of the blade airfoil. Thehigher wedge angles result in leading edge shapes in the outer airfoilsections which improve incidence angle range and operability of thebooster, in addition to being mechanically robust. It may be noted thatthe characteristic of locating Tmax progressively proximate to bladeleading edge in outer span regions, and designing multiple booster rotorstages such that Tmax is located relatively closer to the leading edgein the front stages than the rear stages, as shown for example in FIG.11, are contrary to the conventional practice in the design ofcompression system airfoils. In conventional designs the Tmax locationsof various airfoil sections are chosen based on mechanical designconsiderations such as blade frequencies.

An embodiment of this characteristic of Tmax locations is shown in FIG.11 for the various rotor stages of the booster system. In an example,not meant to be limiting, of this embodiment of the rotor blade, therelative distance is about 0.4 at the root and is about 0.2 at the tip.The variation of the relative distance with respect to the span heightis substantially linear, as shown in FIG. 11. In the preferredembodiment of the booster system, the characteristic variation of therelative distance with span height is substantially the same for therotor blade airfoils in multiple rotor stages, as shown in FIG. 11 forR2, R3, R4 and R5 rotor stages.

One of the ways the operability of the booster system is improved is bydirecting more flow towards the hub region, as the air traverses theaxial path with large curvatures through the booster. One of theparameters of blade design which can be used influence the flowdirections is the dihedral angle at a particular location. Dihedralexists, for example, when the blade surface is not normal to the hub. Asused herein, the definition of “Dihedral” or, alternatively, “DihedralAngle”, is the same as that outlined in the paper “Sweep and DihedralEffects in Axial-Flow Turbomachinery”, Leroy H. Smith, Jr., and HsuanYeh, Journal of Basic Engineering, Transactions of ASME, 62-WA-102,1962.

In another embodiment of a rotor blade, the performance and operabilityof the booster system is improved by adopting a new dihedral angleprofile at the trailing edge 53 that particularly matches the new bladeleading edge 52 sweep rate of change with the span height and thevariation of the location 303 of the maximum airfoil thickness 302described before. FIG. 12 shows an exemplary distribution of thedihedral angle at the trailing edge 53 of the rotor blade with respectto the span height. A negative dihedral angle at a point on the blademeans that the normal to the pressure surface of the blade at thatlocation points towards the longitudinal centerline axis 15 of thebooster system. As shown in FIG. 12, the trailing edge dihedral angle islowest at the blade root 54, adjacent to the booster hub and is negativebetween the blade root 54 and a second height location “H2” 152 (seeFIG. 4) on the trailing edge 53. The dihedral angle becomes lessnegative as the span height increases, becoming positive at anintermediate span height location, thereafter reaching a maximum value,and decreasing thereafter towards the tip.

In yet another embodiment of the rotor blade, the dihedral angle isabout −15 degrees to −20 degrees at the blade root 54, and remainsnegative up to a span height of about 20% from the blade root 54. For anembodiment of a booster system with multiple rotor stages, the trailingedges 53 of the blades in multiple rotor stages have negative dihedralangles near the hub region, from the blade root to about 20% to 30% spanheight.

While there have been described herein what are considered to bepreferred and exemplary embodiments of the present invention, othermodifications of the invention shall be apparent to those skilled in theart from the teachings herein, and it is, therefore, desired to besecured in the appended claims all such modifications as fall within thetrue spirit and scope of the invention.

1. A compression stage for a gas turbine engine comprising: a rotorstage having a plurality of rotor blades spaced circumferentially arounda rotor hub with a longitudinal centerline axis; a stator stage having aplurality of stator vanes spaced in a circumferential direction coaxialwith the longitudinal centerline axis; each stator vane comprising acompressor airfoil having pressure and suction sides extending betweenleading and trailing edges, and longitudinally between a root and a tip,the trailing edge having an exit swirl angle distribution profile fromthe root to the tip such that the exit swirl angle increases between afirst radius location and an intermediate radius location located at aspan height of more than 50% from the root wherein the exit swirl anglehas a maximum value at the intermediate radius location; and each rotorblade having a blade leading edge shape adapted to receive the flow froma stator stage with the exit swirl angle distribution profile.
 2. Acompression stage according to claim 1 wherein the exit swirl decreasesbetween the intermediate radius location and the tip.
 3. A compressionstage according to claim 1 wherein the first radius location is at theroot.
 4. A compression stage according to claim 1 wherein exit swirl atthe tip is more than the exit swirl at the root.
 5. A compression stageaccording to claim 1 wherein the intermediate radius location is locatedat a span height of between about 60% and about 80% from the root.
 6. Acompression stage for a gas turbine engine comprising a rotor stagehaving a plurality of rotor blades spaced circumferentially around arotor hub with a longitudinal centerline axis; a stator stage having aplurality of stator vanes spaced in a circumferential direction coaxialwith the longitudinal centerline axis: each stator vane comprising acompressor airfoil having pressure and suction sides extending betweenleading and trailing edges, and longitudinally between a root and a tip,the trailing edge having an exit swirl angle distribution profile fromthe root to the tip such that the exit swirl angle has a maximum valueat an intermediate radius location between the first radius location andthe tip wherein the intermediate radius location is located at a spanheight of more than 50% from the root; and each rotor blade having ablade leading edge having a sweep angle profile such that the sweepangle increases from the blade root to the blade first height locationat a first rate of change of sweep angle with respect to span heightthat is substantially constant wherein the sweep angle changes from aforward sweep to an aft sweep at a location that is between 10 percentand 30 percent of the span.
 7. A compression stage according to claim 6wherein the first rate of change of sweep angle with respect to spanheight is less than the second rate of change of sweep angle withrespect to span height.
 8. A compression stage according to claim 6wherein the first rate of change of sweep angle with respect to spanheight is substantially equal to the second rate of change of sweepangle with respect to span height.
 9. A booster for a gas turbine enginecomprising: an inlet guide vane stage having a plurality of inlet guidevanes spaced in a circumferential direction around a longitudinalcenterline axis; a plurality of stator stages, each stator stage havinga plurality of stator vanes spaced in a circumferential direction aroundthe longitudinal centerline axis and each stator vane having pressureand suction sides extending between leading and trailing edges, andlongitudinally between a root and a tip, the leading edge for each vanein at least one stator stage having a forward sweep from the root to afirst span location and at least one stator stage having vanes havingleading edge aft sweep between a first span location and the tip, theleading edge aft sweep at the tip for each stator vane with leading edgeaft sweep is less than the leading edge aft sweep at the tip for anyvane located axially forward from it; a plurality of rotor stages, eachrotor stage having a plurality of rotor blades spaced in acircumferential direction around a rotor hub located coaxially aroundthe longitudinal centerline axis, each rotor blade having a bladeleading edge shape adapted to receive the flow from a stator stage or aninlet guide vane stage located axially forward from it.
 10. A boosteraccording to claim 9 wherein at least one stator stage has stator vaneshaving leading edge forward sweep at the root and leading edge aft sweepat the tip.
 11. A booster according to claim 9 wherein the leading edgeaft sweep for the vanes in at least one stator stage increases betweenthe first radius and the tip.
 12. A booster according to claim 9 whereinall of the vanes have a leading edge forward sweep between the root andan outer radius location on the vane.
 13. A booster for a gas turbineengine comprising: an inlet guide vane stage having a plurality of inletguide vanes spaced in a circumferential direction around a longitudinalcenterline axis; a plurality of stator stages, each stator stage havinga plurality of stator vanes spaced in a circumferential direction aroundthe longitudinal centerline axis and each stator vane having pressureand suction sides extending between leading and trailing edges, andlongitudinally between a root and a tip, the leading edge for each vanein at least one stator stage having a forward sweep from the root to afirst span location and at least one stator stage having vanes havingleading edge aft sweep between the first span location and the tip, theleading edge aft sweep at the tip for each stator vane with leading edgeaft sweep is less than the leading edge at sweep at the tip for any vanelocated axially forward from it; and a plurality of rotor stages, eachrotor stage located in the upstream direction from a stator stage andcoaxial with the stator stage and having a plurality of rotor bladesspaced in a circumferential direction around a rotor hub locatedcoaxially around the longitudinal centerline axis, each rotor bladehaving a blade leading edge having a sweep angle that increases from theblade root to a blade first height location on the blade leading edge ata first rate of change of sweep angle with respect to the span heightthat is substantially constant and increases from the blade first heightlocation to the blade tip of the rotor blade at a second rate of changeof sweep angle with respect to the span height that is substantiallyconstant wherein the sweep angle changes from a forward sweep to an aftsweep at a location that is between 10 percent and 30 percent of thespan.
 14. A booster according to claim 13 wherein the first rate ofchange of sweep angle with respect to span height is less than thesecond rate of change of sweep angle with respect to span height.
 15. Acompression stage according to claim 13 wherein the first rate of changeof sweep angle with respect to span height is substantially equal to thesecond rate of change of sweep angle with respect to span height.
 16. Acompression stage according to claim 6 wherein the intermediate radiuslocation is located at a span height of between about 60% and about 80%from the root.